This invention is particularly concerned with military aircraft as opposed to civilian or commercial aircraft and even more particularly to the class of aircraft that would fall in the fighter class. Fighter aircraft are typically powered by after-burning gas turbine engines that produce thrust by compressing air in a compressor section, adding fuel to the air and burning it in a combustor section to produce combustion products. The combustion products then flow through a turbine section which drives the compressor section, and the combustion products then exit the engine through a variable area, convergent/divergent exhaust nozzle. To boost the engine's thrust for brief periods, most fighter aircraft engines incorporate an augmentor in an augmentor duct located between the turbine section and the exhaust nozzle. When maximum thrust is desired, large amounts of fuel are fed into the augmentor duct and burned immediately downstream of the engine's turbines section. Although this "after-burning" produces a desirable increase in thrust, it also exposes the exhaust nozzle, and particularly the flaps and seals in the convergent section of the exhaust nozzle, to combustion gas that is at temperatures well beyond that necessary to burn through the flaps and seals.
To prevent such burn-through, cooling air is provided to the flaps and seals to maintain their temperature at an acceptable level. This cooling air flows to the flaps and seals from an annular cavity between the augmentor duct and the augmentor liner to another cavity between the flaps and seals and the nozzle support structure radially outward therefrom. Although this type of cooling scheme has proven to be quite effective, burn-through of the flaps or seals during augmentor operation can still occur.
Usually, a burn-through at the nozzle flaps or seals results from a loss of cooling air. This loss in cooling air leads to a hardware anomaly, loss of combustion gas containment, increase in nozzle hardware temperature and subsequent fire. Since propagation of the fire proceeds at an extremely rapid rate until after-burning is terminated, response time is critical.
Currently, detection of a nozzle burn through occurs by an airfield tower report, a wingman report, a noticeable loss in thrust, or by burn-through detector systems mounted on the aircraft's airframe. Unfortunately, by the time any of these detection methods notifies the pilot of the burn-through, a significant safety risk is present. In addition, considerable engine/airframe hardware damage is likely to have occurred before the pilot terminates afterburning.
What is needed is an integral gas turbine engine augmentor nozzle burn-through detection system that reduces pilot response time compared to current methods.